1. Field of the Invention
This invention pertains to gas turbine engines and pertains more particularly to improvements in pickup and delivery of cooling airflow with minimum static pressure loss therein.
2. Description of the Prior Art
High performance gas turbine engines utilize a high speed compressor section to deliver pressurized airflow to the engine combustor. High energy exhaust gases from the combustor drive one or more high temperature turbines, which, in turn provide the motive power for the compressor section. It is conventional to use a small portion of pressurized air from the compressor section for cooling of internal hot section components of the engine as well for seal buffering. This parasitic cooling flow, sometimes referred to as a secondary bleed airflow, represents a parasitic power loss of the engine, and it is therefore important that such cooling flow be minimized. Thus, this cooling flow should not be subject to undue pressure losses nor unnecessary heating.
The flow path for this cooling flow in many instances must radially inwardly traverse the rotating shafting arrangement which is driving the compressor in order to reach a radially inner location for subsequent delivery in axial directions to components to be cooled or seals to be buffered. In certain gas turbine engine designs it is highly desirable or even a requirement that the shafting arrangement which is transmitting torque to drive the compressor stage be at a location which has a much larger radius than the smaller radius to which the cooling airflow needs to be delivered. In passing to such a smaller radius the cooling airflow will naturally experience an increase in tangential velocity, introducing a vortex action into the cooling airflow. This introduces significant pressure drop to the cooling airflow as it passes to the smaller radius.
Loos U.S. Pat. No. 2,618,433 represents a bleed airflow arrangement suitable only to much older gas turbine engine technology which was constrained, as mentioned in the Loos patent, by the lack of incorporation of a diffuser in the compressor section. This problem has been overcome in newer gas turbine engines. As a result, Loos teaches inclusion of bleed air passageways opening directly into the compressor stages immediately down stream of a set of compressor blades such that highly swirled, vortically spinning air is bled from the compressor. Loos then teaches utilization of a blading structure in order to recover work from the swirling flow to assist in driving the compressor section.
Rieck U.S. Pat. No. 4,541,774 represents a more modem gas turbine engine and cooling airflow pickup system. The arrangement illustrated in the Rieck patent is of the type wherein the shafting 24 and torque transmittal path for the compressor section is very nearly at the same radius as the internal radius to which the cooling airflow must be delivered--as represented by the small radial difference between holes 32 and space 18 after the rotor 22. Thus, after radially inwardly traversing shaft 24 the cooling airflow is already essentially at its required radial location and thus not subject to further increase in vortex spinning thereof due to movement to a much smaller radial location. The Rieck reference also represents a certain form of tangential onboard injection systems (TOBI) of many gas turbine. Specifically, like other TOBI systems, the Rieck reference utilizes a set of stationary turning vanes 34 which deliberately induces rotational vortex flow into the cooling airflow at the expense of reduced pressure of the cooling airflow.